1. Field of the Invention
This present invention relates to turbine components equipped in gas turbine engines for aircraft, more particularly, to turbine components equipped in cooled turbine blades cooled with a film cooling.
2. Description of the Related Art
A conventional cooled turbine blade as a cooling turbine component in a gas turbine engine will be described below.
A conventional cooled turbine blade comprises a blade body (a main body), and inside the blade body it includes a cooling passage. Compressed air is fed from a compressor or a fan, and this air provides cooling air to the cooling passage.
Film cooling holes are provided on the surface of the blade body by an electric discharging or laser machining. Each film cooling hole is connected to the cooling passage. The film cooling holes further include an inlet opening which introduces the cooling air and is provided on an inlet profile of the blade body and a wedge shaped spurt out opening which spurts out the cooling air and is provided on an outlet profile of the blade body. Furthermore, each film cooling hole is inclined to a thickness direction of the blade body, so that the cooling air coats the blade body profile.
Accordingly, when the gas turbine engine is operating, the compressed air is bled from the compressor or the fan, and some of the compressed air (the cooling air) is introduced to the cooling passage. The cooling air is introduced to the inlet opening and spurted from the spurt out opening, thereby forming a cooling film which covers and protects the blade profile. According to the structure mentioned above, the cooled turbine blade is cooled with a film cooling as well as a convection cooling which cools the inside of the blade.
Meanwhile, an invention that is described in Japanese Patent Application Laid-Open No. 2002-221005 shows one of the conventional art.
Heretofore, the conventional cooling hole is formed in the diffuser which is shaped in such that the outlet side is expanded in an expansion angle, so that the cooling air uniformly coats the blade profile. With the angle of the spurt out opening set sizably, the number of the film cooling holes necessary for coating the blade body profile can be reduced. In this case, however, since the amount of cooling air which is fed from the fan or the compressor is constant, the current velocity of the cooling air at the spurt out opening becomes extremely slow, the separation phenomenon occurs, and the cooling performance of the cooled turbine blade is also reduced. In addressing these problems, conventionally to obtain a desired cooling performance, a lot of film cooling holes are provided in the blade profile instead of making the expansion angle larger. In this case, however, longer manufacturing time is needed for making the cooled turbine blade, and the production costs are increased as well. Additionally, as the amount of cooling air necessary for a desired cooling performance is increased, the engine efficiency of the gas turbine engine is lowered.